Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose blade assemblies to these high temperatures. As a result, blades must be made of materials capable of withstanding such high temperatures. Blades and other components often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Blades typically extend radially from a rotor assembly and terminate at a tip within close proximity of the blade rings. The blade rings may be exposed to the hot combustion gases and, similar to the blades, the blade rings often rely on internal cooling systems to reduce stress and increase the life cycle. The blade rings are spaced radially from the blade tips to create a gap therebetween to prevent contact of the blade tips with the blade rings as a result of thermal expansion of the blades. During conventional startup processes in which a turbine engine is brought from a stopped condition to a steady state operating condition, blades and blade rings pass through a pinch point at which the gap between the blade tips and the blade rings is at a minimal distance due to thermal expansion. The blade tips of many conventional configurations contact or nearly contact the blade rings. Contact of the blade tips may cause damage to the blades. Furthermore, designing the gap between the blade tips and the blade rings for the pinch point often results in a gap at steady state conditions that is larger than desired because the gap and combustion gases flowing therethrough adversely affect performance and efficiency.
FIG. 1 shows a cross-section through a portion of a turbine engine. A turbine engine 10 can generally include a compressor section 12, a combustor section 14 and a turbine section 16. A centrally disposed rotor 18 can extend through the three sections.
Generally, the combustor section 14 is enclosed within a casing 20 that can form a chamber 22, together with the aft end of the compressor casing 24 and a housing 26 that surrounds a portion of the rotor 18. A plurality of combustors 28 and ducts 30 can be provided within the chamber 22, such as in an annular array about the rotor 18. Each duct 30 can connect one of the combustors 28 to the turbine section 16.
The turbine section 16 can include an outer casing 32 that encloses alternating rows of stationary airfoils 34 (commonly referred to as vanes) and rotating airfoils 36 (commonly referred to as blades). Each row of blades can include a plurality of airfoils 36 attached to a disc 38 provided on the rotor 18. The rotor 18 can include a plurality of axially-spaced discs 38. The blades 36 can extend radially outward from the discs 38 and terminate in a region known as the blade tip 40.
Each row of vanes can be formed by attaching a plurality of airfoils 34 to the stationary support structure in the turbine section 16. For instance, the airfoils 34 can be supported by a vane carrier 42 that is attached to the outer casing 32. The vanes 34 can extend radially inward from the vane carrier 42 or other stationary support structure to which they are attached.
In operation, the compressor section 12 can compress ambient air. The compressed air 44 from the compressor section 12 can enter the chamber 22 and can then be distributed to each of the combustors 28. In the combustors 28, the compressed air can be mixed with the fuel introduced through a fuel nozzle 46. The air-fuel mixture can be burned, thereby forming a hot working gas 48. The hot gas 48 can flow through the ducts 30 and then through the rows of stationary airfoils 34 and rotating airfoils 36 in the turbine section 16, where the gas 48 can expand and generate power that can drive the rotor 18. The expanded gas 50 can then be exhausted from the turbine 16.
It should be noted that each row of blades 36 is surrounded by the stationary support structure of the turbine, which can be the outer casing 32, the vane carrier 42 or a ring seal (not shown). The space between the blade tips 40 and the neighboring stationary structure is referred to as the blade tip clearance C. During engine operation, gas leakage can occur through the blade tip clearances C, resulting in measurable engine performance decreases in power and efficiency.
The compressed air 44 from the compressor 12 can be used to cool the rotor 18 or to internally cool the turbine blades 36, among other things. A portion 52 of the compressed air 44 from the compressor 12 can be extracted from the chamber 22 and routed externally of the engine 10 through a fluid conduit connected in fluid communication with the chamber 22. The fluid conduit can be a single conduit or a plurality of conduit segments. The fluid conduit includes a first conduit segment 54 and a second conduit segment 60. By entering the first conduit segment 54, the portion of air 52 bypasses the combustors 28. The portion of air 52 is cooled by an external cooler 56 disposed along the fluid conduit. However, gas turbines undergo a range of operation for which different blade clearances are required. Cooling systems that do not account for the range of operation of the gas turbine and the change of clearances from start up through steady state operation impart inefficiencies into the system.
While small blade tip clearances C are desired to minimize gas leakage, it is critical to maintain a clearance C between the rotating turbine components (blades 36, rotor 18, and discs 38) and the stationary turbine components (vanes 34, outer casing 32, vane carriers 42 and ring seals) at all times. Rubbing of any of the rotating and stationary components can lead to substantial component damage, performance degradation, and extended outages.
However, during transient conditions such as during engine startup or part load operation, it can be difficult to ensure that adequate blade tip clearances C are maintained because the rotating parts and the stationary parts thermally expand at different rates. For instance, in a cold start situation, the rate of thermal expansion of the thermal stationary support structure is at least initially less than the rate of thermal expansion of the rotating turbine components due to the relatively larger size and thickness of the stationary support structure. As a result, the blade tip clearances C can actually decrease because the rotating components expand radially outward faster than the stationary support structure, raising concerns of blade tip rubbing.
To avoid blade tip rubbing, large tip clearances are initially provided so that minimum blade tip clearances C are maintained at known pinch points, that is, during operational conditions where the clearances C would otherwise be expected to be the smallest (hot restart, spin cool, etc.). However, because the minimum blade tip clearances C are sized for these pinch point conditions, the clearances C eventually become overly large as the rate of thermal expansion of the rotating components slows or substantially stops while the stationary support structure continues to grow radially outward. Such oversized clearances C can occur as the engine approaches or attains steady state operation, such as at base load. Consequently, engine power and efficiency can be reduced.
Thus, there is a need for a system that can improve engine performance by reducing or minimizing blade tip clearances at desired engine operating conditions.